Gas turbines are known to comprise the following elements: a compressor for compressing air; a combustor for producing a hot gas by burning fuel in the presence of the compressed air produced by the compressor; and a turbine for expanding the hot gas produced by the combustor. Gas turbines are known to emit undesirable oxides of nitrogen (NO.sub.x) and carbon monoxide (CO). One factor known to affect NO.sub.x emission is combustion temperature. The amount of NO.sub.x emitted is reduced as the combustion temperature is lowered. However, higher combustion temperatures are desirable to obtain higher efficiency and CO oxidation.
Two-stage combustion systems have been developed that provide efficient combustion and reduced NO.sub.x emissions. In a two-stage combustion system, diffusion combustion is performed at the first stage for obtaining ignition and flame stability. Premixed combustion is performed at the second stage to reduce NO.sub.x emissions.
The first stage, referred to hereinafter as the "pilot" stage, is normally a diffusion-type burner and is, therefore, a significant contributor of NO.sub.x emissions even though the percentage of fuel supplied to the pilot is comparatively quite small (often less than 10% of the total fuel supplied to the combustor). The pilot flame has thus been known to limit the amount of NO.sub.x reduction that could be achieved with this type of combustor. In a diffusion combustor, the fuel and air are mixed in the same chamber in which combustion occurs (i.e., a combustion chamber).
Pending U.S. patent application Ser. No. 08/759,395, assigned to the same assignee hereunder (the '395 application), discloses a typical prior art gas turbine combustor. As shown in FIG. 1 herein, combustor 100 comprises a nozzle housing 6 having a nozzle housing base 5. A diffusion fuel pilot nozzle 1, having a pilot fuel injection port 4, extends through nozzle housing 6 and is attached to nozzle housing base 5. A plurality of main nozzles 2, each having at least one main fuel injection port 3, extend substantially parallel to pilot nozzle 1 through nozzle housing 6 and are attached to nozzle housing base 5. Fuel inlets 16 provide fuel 102 to main nozzles 2. A main combustion zone 9 is formed within a liner 19. A pilot cone 20, having a diverged end 22, projects from the vicinity of pilot fuel injection port 4 of pilot nozzle 1. A pilot flame zone 23 is formed within pilot cone 20 adjacent to main combustion zone 9.
Compressed air 101 from compressor 50 flows between support ribs 7 through main fuel mixers 8. Each main fuel mixer 8 is substantially parallel to pilot nozzle 1 and adjacent to main combustion zone 9. Within each main fuel mixer 8, a plurality of flow turbulators 80, such as swirler vanes, generate air turbulence upstream of main fuel injection ports 3 to mix compressed air 101 with fuel 102 to form a fuel/air mixture 103. Fuel/air mixture 103 is carried into main combustion zone 9 where it combusts. Compressed air 101 also enters pilot flame zone 23 through a set of stationary turning vanes 10 located inside pilot swirler 11. Compressed air 101 mixes with pilot fuel 30 within pilot cone 20 and combusts in pilot flame zone 23.
FIG. 2A shows a radial cross-sectional view of prior art gas turbine combustor 100 taken along line A--A thereof. As shown in FIG. 2A, pilot nozzle 1 is surrounded by a plurality of main nozzles 2. Pilot swirler 11 surrounds pilot nozzle 1. A main fuel mixer 8 surrounds each main nozzle 2. Main fuel mixers 8 are separated from one another by a distance, d. In the embodiment shown in FIG. 2A, main fuel nozzles 2 are disposed uniformly around pilot nozzle 1. Consequently, distance, d, between adjacent main fuel mixers 8 is nearly the same for each pair of adjacent main fuel mixers 8, although it may be variable. Fuel/air mixture 103 flows through main fuel mixers 8 (out of the page) into main combustion zone 9 (not shown in FIG. 2A). Pilot swirler 11 forms an annulus 18 with liner 19. Compressed air 101 flows through annulus 18 (out of the page) into main combustion zone 9. Note that compressed air 101 flowing through annulus 18 is not premixed with any fuel.
FIG. 2B shows a radial cross-sectional view of prior art gas turbine combustor 100 taken along line B--B thereof. As shown in FIG. 1, line B--B is downstream of line A--A. Line B--B is adjacent to main combustion zone 9, downstream of main nozzles 2 and pilot nozzle 1. As shown in FIG. 2B, a plurality of main fuel mixers 8 are disposed uniformly around pilot swirler 11. Pilot swirler 11 forms an annulus 18 with liner 19. Compressed air 101 flows through annulus 18 (out of the page) into main combustion zone 9. Note that compressed air 101 in annulus 18 is not premixed with any fuel.
As shown in FIG. 2B, main fuel mixers 8 are separated from one another by distance, d. Although, as described above, distance, d, between adjacent main fuel mixers 8 may be variable or nearly constant, it is important to note that the distance between a given pair of main fuel mixers in FIG. 2B is substantially the same as the distance between the same pair of main fuel mixers 8 as shown in FIG. 2A. Thus, each main fuel mixer 8 is separated from every other main fuel mixer 8 and each main fuel mixer 8 is nearly constant in cross-sectional area along its length.
While gas turbine combustors such as the combustor disclosed in the '395 application have been developed to reduce NO.sub.x and CO emissions, current environmental concerns demand even greater reductions. It is known that leaner, more homogeneous fuel/air mixtures burn cooler and more evenly, thus decreasing NO.sub.x and Co emissions. Since, in a premix combustor, main stage fuel and compressed air are mixed in main stage fuel mixers before combustion occurs, there is a need in the art for a main stage fuel mixer that reduces NO.sub.x and CO emissions from gas turbine combustors by providing leaner, more homogeneous fuel/air mixtures for main stage combustion.